SpaceX advances drive for Mars rocket via Raptor power

SpaceX Co-Founder and Vice President of Propulsion Development Tom Mueller has revealed the company is deep into the development of the first “full flow methane-liquid oxygen” rocket engine. Known as the Raptor, nine of these immensely powerful engines – on one or three cores – will be utilized to send SpaceX’s Super Heavy Lift Launch Vehicle (SHLV) uphill on missions to Mars.

The Raptor Engine History:

SpaceX has proven to be extremely innovative in their production of engines – particularly aggressive on development and manufacturing cost. However, their technology, though innovative, has rarely stepped into the unproven.

As advanced as the Merlin 1D is – having the highest thrust/weight ratio in commercial use, being the most efficient RP-1 gas generator ever developed in the US, and probably one of the best thrust/cost relationships in the world – it is based on the most basic rocket cycle for a turbopump fed system, the gas generator.

In fact, the company has always deployed the most simple cycles and propellants in its rocket engines and thrusters. Not a bad choice given they can beat even the Chinese on price and have shown, up to now, good reliability.

However, for the more demanding missions, a higher performance technological approach might be required, particularly for tackling a problem as demanding as a fully reusable Mars colonization architecture.

At the AIAA Joint Propulsion conference on July 30, 2010 then SpaceX McGregor rocket development facility director Tom Markusic provided information from the initial stages of planning for two families of dual stage exploration-class launchers and two new rocket engines to power them.

The first engine presented was a kerosene and liquid oxygen powered engine, of the gas generator cycle, that was even more powerful than the mighty F-1 that powered the Saturn V. This engine would power both first stage of the Falcon X and Falcon XX.

The Merlin 2 engine, would have been capable of a projected 7,600 kN (1,700,000 lbf) of thrust at sea level and 8,500 kN (1,920,000 lbf) in a vacuum. The engine would have been the most efficient of its kind, with specific impulse (a measure of the efficiency of propellant usage) being even better than the current Merlin 1D.

However, the most interesting engine shown was the Raptor engine.

In a complete break with the company’s tradition, it introduced both a new propellant and a new engine cycle.

The Raptor engine, as presented, was planned as a staged combustion, liquid hydrogen and oxygen engine with a vacuum thrust of 150klbf (667kN) and 470s of isp – designated to power the upper stage of the super heavy rockets.

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This development path was unexpected from a company solely focused on low cost, since it would require expertise in the handling of cryogenic liquid hydrogen, which at −423.17 °F (20.28 K) is far colder than either liquid nitrogen (−321 °F or 77 K), liquid oxygen (−297.33 °F or 90.19 K) or liquid methane (−258.68 °F or 111.66 K).

Hydrogen also has a tendency to weaken metals by embrittlement, and – as a propellant – it requires three times the tank’s volume of the RP-1/LOX that the company has been using since their beginnings with the Falcon 1. Not to mention the fact of having to implement a dual fuel ground supply and support equipment at the pad, would significantly increase operating costs and expenses. Exactly the sort of cost structure the company’s founder, Elon Musk, had repeatedly stated he wanted to avoid.

Yet, later that year, Mr. Musk stated publicly that the presentation was based on “brainstorming ideas” and a “bunch of ideas for discussion.” Mr. Markusic soon left the company, a move that generated much speculation on the disclosed material.

The first clues that SpaceX was seriously considering the use of staged combustion methane engines came in a series of questions asked with respect to a Request For Information of the Air Force for the Reusable Booster System High Thrust Main Engine during May, 2011.

The request specified a set of performance requirements that could only by covered by very high performance staged combustion engine, like the AJ-26-500 or the RD-191. The information specified kerosene and liquid oxygen as the propellant of choice.

SpaceX asked if the Planning Directorate was interested in engines they were designing and planning to build for other customers, and whether a response that offered both RP and methane propellant – possibly with fly-off testing – would be an acceptable response.

Since the specified thrust was 300~500klbf (1,300~2,200kN) and the minimum sea level isp 300s, it implied a completely different engine from the Merlin 1 family.

In later conferences, specially through 2012 and 2013, Mr. Musk noted the future propellent they would use for their plans of Mars colonization would be liquid methane with liquid oxygen.

This was in related to the fact it could be sourced from Mars, was easier to store and handle than liquid hydrogen and at just a 27% more volumetric than the RP-1/LOX combination, the tank and T/W of the engine could be kept at optimal sizes. He also stated that they would start using the staged combustion engine cycle.

Additionally, Mr. Musk also introduced the mysterious MCT project, which he later revealed to be an acronym for Mars Colonial Transport. This system would be capable of transporting 100 colonists at a time to Mars, and would be fully reusable.

Speculation noted the comparison to the 2010 Russian Reusable Launch Vehicle, or MRKN, study – which had settled on 450klbf (2,000KN) of thrust, and traded a kerosene and a methane version of staged combustion engines from each of the two main propulsion companies in Russia, namely NPO Energomash’s RD-191 (kerosen) and RD-192 (methane), and the KbKhA’s RD-0162 (methane) and RD-0163 (kerosene).

The first conclusion from the evaluations was that methane was the best propellant choice.

As per most trades between kerosene and methane, the performance was basically the same, albeit with a slight advantage for kerosene, which ended up being lighter and a bit smaller. However, methane was found to be more efficient.

Methane also holds much better reusability properties on the engine, whereas Kerosene tends to polymerize (coke) and thus requires oxidizer rich combustion, which is quite corrosive and aggressive to the turbopump system. Kerosene also leaves more residue throughout the engine, which might require expensive cleaning and even rebuilding.

Methane can be run by preburning the fuel, has basically no coking problems, has much better cooling characteristic, and if used in the form of LNG, is the cheapest and most abundant fuel.

The main source of deliberation was that the RD-0162 could actually be run at 133%, providing an effective thrust of 600klbf. As such, it appeared that both the Russians and SpaceX had reached very similar propulsion solutions.

Mr. Mueller confirmed nine of these engines would power each 10 meter diameter core of the notional MCT. “I’m quite proud to have my name attached to this engine,” said the SpaceX Co-Founder at the event.

The implications of this revelation are numerous – the most important being SpaceX are now fully treading into uncharted territory.

There have been just two full flow projects that actually hit a test stand: the Glushko’s RD-270 and the Rocketdyne Aerojet Integrated Powerhead Demonstrator.

In a sense, Raptor appears to share some the objectives with each of those projects.

The Previous Full Flow Engines:

During the Moon Race, the N-1 and OKB-1 (now RSC Energia) Chief Designer Sergei Korolev had a serious disagreement with the OKB-456 (now NPO Energomash) Chief Designer Valentin Glushko, who had, up to that point, supplied all the main engines for its first stages.

Officially Korolev’s Chief Deputy, Vasily Mishin, had asked for way too advanced specifications on the engine requirements and insisted on kerosene or hydrogen and liquid oxygen as propellant.

Glushko had offered a most advanced and powerful engine, but only with the highly toxic hypergolic propellent combination, in which he had a lot of experience, specially on the staged combustion cycle. The discussion escalated to a closed door shouting match between Korolev and Glushko. They never talked again.

Korolev handed over the task of designing N-1 engines to the aircraft turbine manufacturer Kutznesov and Glushko sided with Korolev’s opponent Vladimir Chelomei, Chief Designer at OKB-52. They came up with the UR-700 project for the Moon race.

The UR-700M, a 35,000,000lb (16,000 tonnes) monster rocket would have dwarfed even the Saturn V and would have been the rocket to enable the Soviets to conquer Mars. A similar destiny that SpaceX sees on the Raptor.

To power such a project, Glushko decided to use the most advanced cycle for turbopump fed engines, the full flow or full staged combustion. It used the hypergolic combination of N2O4 and UDMH as propellant.

Such an engine, with a single nozzle RD-270, had a sea level thrust of 1,400klbf (6.3MN) and an isp of 301s, while in a vacuum it provided a thrust of 1,500klbf (6.7MN) and an isp of 322s.

Not only was this the most powerful “per nozzle” engine ever attempted in the USSR, but it had an amazing 127 T/W ratio at sea level and sported an unheard of 3,858psi (26.6MPa) pressure in the main combustion chamber. A record that not even NPO Energomash’s latest RD-191, at just 3,727psi (25.7MPa), could match.

From October 23, 1967 to July 24, 1969 this engine hit the test stand and 22 prototypes performed a total of 27 firing. Only nine of those tests were nominal. While the most difficult problems were overcome, instability problems where not completely solved and the project was axed as part of the UR-700 project cancellation.

However, to this day, the Russian engine that was destined to enable Mars exploration still holds the biggest thrust per nozzle record for any staged or full staged combustion engine.

The second project was the 250klbf (1,100kN) Integrated Powerhead Demonstrator, part of the joint DoD/NASA Integrated High Payoff Rocket Propulsion Technologies (IHPRPT) program.

The objective of this program was to test high payoff technologies that would enable higher performance than the SSME, but with a useful life of up to 200 missions. In keeping with the USA engine expertise, it used liquid hydrogen and oxygen as propellant.

The main contractor was Rocketdyne, but Aerojet was in charge of the critical oxygen rich preburner and the channel wall nozzle. The IPD project also looked into developing hydrostatic bearings.

Current turbopumps rotate on balls or rolling elements, with the rotation producing heat and wear and tear due to material contact. Hydrostatic bearings use the turbopump’s own fluid to actually float the pump on high pressure liquid.

The advantage is the reduction of wear being suffered only during start up, in turn significantly increasing useful life and enabling very long mission duration. However, if you run the turbine with hot gas of the other propellant element (i.e. oxidizer gas with fuel liquid or the other way around), you have a perfectly explosive device.

Thus, this technology is best used on engines that have completely separate oxidizer and fuel turbopump systems, in turn eliminating a failure mode by not requiring an interseal.

In gas generator and staged combustion systems, a seal must separate the fuel side from the oxidizer side of the turbopump. Any seal failure would be catastrophic, and using separate systems completely eliminates the risk.

The breakthrough characteristic is that since each turbine is effectively fed by its own propellant mass, it can have a lot more turbine power.

Such power could be used to increase the main combustion chamber pressure and increase the overall performance, or by using cooler gases, providing the same performance as a staged combustion engine but with much less stress on materials and thus significantly reduce material fatigue or weight.

As an added bonus, lower pressures are required through the pumping system, which not only increases the life span, but reduces the risk and effects of a catastrophic failure.

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